Theoretical and experimental data for a number of NACA 6A-series airfoil sections

THEORETICAL AND EXPERIMENTAL DATA FOR A NUMBER OF NACA 6A--SERIES AIRFOIL SECTIONS

TABLE III.--OR.DINATES OF NACA 64A010 AIRFOIL

TABLE I.--ORDINATES OF NACA 63A010 AIRFOIL

SECTION

SECTION [Stations and ordinates given in percent of airfoil chord] ]stations and ordinates given in percent of airfoil chord] Upper surface Lower surface ill)per surface Lower surface St,_.tion Ordinate Ordinate Station Station Ordinate Station Ordinate 0 0 0 0 O 0 0 .5 .804 .5 --. 804 .5 • 810 .5 --.816 --. 983 .75 • 969 .75 --. 969 • 75 .983 .75 -- I. 225 1.25 1.250 1.25 --1.2,50 1.25 1.225 1•25 1.688 2.5 -- 1.688 2.5 1. 737 2. 5 --1.7:37 2.5 --2. 412 5.0 2.327 5.0 --2. 327 5.0 2. 412 5. 0 7.5 2.917 7.5 --2. 917 7.5 2. 805 7. 5 --2. 805 l0 3. 199 10 --3. 199 10 :3. 324 10 --3. 324 --3. 950 15 3.813 15 --3.813 15 3. 950 15 --4. 272 2O ,L 4(XI 20 -- 4.49O 20 4. 272 20 25 4. 606 25 --4. 606 25 4. 714 25 --4. 714 30 --4. 913 30 4. 837 30 --4. 837 4. 913 :_ --4.99S 35 4. 995 35 --4. 995 35 4. 968 35 40 4. 995 40 --4. 995 4O -L 968 40 --4. 968 --4. 837 45 4. 894 45 --4. 894 45 4. 837 45 50 4. 613 f_l --4. 613 50 4. 684 50 --4. 684 --4. 388 55 4.3tl 55 --4.311 55 4. 388 55 60 4. 021 60 --4. 021 60 3. 943 60 --3. 943 65 3. 517 t{5 --3. 517 65 3. 597 65 --3. 597 --3.127 70 3. 044 70 -- 3. 044 70 3.127 70 75 2. 623 75 --2. 623 75 2. 545 75 --2. 545 2. 040 80 --2.040 8O 2.103 80 --2. 103 8O --1.582 85 I, 535 85 -- i. 535 S5 1..582 85 90 l. 062 90 --1.062 930 1• 030 (30 --1.030 • 525 95 --. 525 95 .541 95 --. 54I 100 • 021 100 --. 021 19O • 02t I(X) --. 02l L. E. radius: 0.742 L. E. radius: 0.687 T. E. radius: 0.023 T. E. radus: 0.023

TABLE IV.--ORDINATES OF NACA 64A210 AIRFOIL

TABLE II.--ORDINATES OF NACA 63A210 A[RFOIL

SECTION SECTION

]Stations and ordinates given in percent of airfoil chord] ]Stations and ordinates given in percent of airfoil chord] Upper surface Lower surface Upper surface Lower snrface Station Ordinate Ordinate Station Station Ordinate Ordinate Station 0 0 0 O 0 0 0 0 .856 .576 --. 744 • 423 .868 .577 -. 756 .424 .664 --.900 •665 1.044 .835 --. 886 1. 058 .836 1.151 1. 367 1. 349 --1. 125 1.153 1. 342 1. 347 --1.19O 1.895 2. 613 2. 384 1. 944 2. 616 -- 1. 522 2•387 -- 1.473 4. 869 2. 769 5. 131 --2. 047 4.874 2.685 5. t26 -- 1.963 7. 364 3. 400 7.6.36 --2. 428 7.369 3.288 7.63l --2. 316 9.868 3.792 I0.132 --2. 600 9. 863 3. 917 10. t37 --2. 725 14. 869 4. 729 15.131 --3. 167 14.874 4.592 15.126 --3. 030 19. 882 5.328 20. I18 --3. 4(_ 19.885 5.200 20.115 --3. 340 24.9{_1 5. 656 25. 100 24. 89!/ 5. 764 25.102 --3. fi62 --3. 554 29. 916 6. 06O 30. 084 --3. 761 29.917 5.984 30.083 --3. 688 34. 935 0. 219 35. 065 --3. 771 34.935 6.192 35.065 --3. 744 39.955 6.274 40.045 --3. 716 39, 955 6. 247 40. 045 --3. 689 44. 975 6. 151 45. 025 --3. 523 44.975 6.208 45.025 --3. 580 49. 994 5. 943 _. 006 --3.28:{ 49.994 6.014 50.000 -3. 354 55.012 5. 637 54. 988 -- 2. 985 55.012 5. 714 54.988 --3. O62 --2. 641 60.028 5.323 59.972 --2. 710 60. {)28 5. 245 59. 972 65. 04i 4. 772 64. 959 --2. 262 65.042 4. 852 64. 958 --2. 342 70.052 4. 227 69. 948 -- l, 86l 70.054 4. 31[) 69. 946 -- I. 944 3. 624 74. 939 -- 1. 464 75.063 3.702 74.937 -- 1. 542 75. 061 80. 074 2. 974 79. 926 -- 1. 104 8{1.076 3.{137 79.924 --1. 167 85.072 2. 254 84. 928 --. 812 85.074 2.301 84.926 --. 859 90.052 • 1.551 89.948 --. 571 (30.050 1. 519 89. 950 --. 539 95. 026 • 769 94. 974 --. 279 95.027 .785 94.974 --. 295 i ' i 10{). [g)0 .021 1O0. (}_) --. 021 100. 0_) .021 100.9O0 --. 021 L.E. radius: 0.742 L. E. radius: 0•687 i T. E. radius: 0.023 T. E. radius: 0.023 Slope of radius through L. E.: 0.095 i " Slope of radius thrtalgh L. E.: 0.095 l I

I0 REPORT NO. 903--NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

TABLE VII.--ORDINATES OF NACA 642A215 AIRFOIL

TABLE V.--ORDINATES OF NACA 64A410 AIRFOIL

SECTION

SECTION

[Stations and ordinates given in percent of airfoil chordl [Stations and ordinates given in percent of airfoil chord] Upper surface Lower surface Upper surface Lower surface Station Ordinate Stat ion Ordinate Station Ordinate Station Ordinate O o 0 0 0 O 0 --1.131 .388 1. 243 .612 .350 .902 .050 --. 078 1. 509 .876 --1.351 --. 796 .624 .582 1.112 .918 1.930 1.393 -1.688 1. 451 1. 441 --. 969 1.107 1.059 --2. 29l 2.333 2. 713 2. 667 2.276 2.695 2.724 -- 1. 251 --3.111 --1.592 4.811 3. 833 5. ! 80 4.749 3.034 5.251 4. 683 7. 690 --3. 711 3. 865 7. 770 --1.919 7.304 7. 230 5. 391 10. 198 --4.199 9.737 4.380 10.263 --1.996 9. 802 --4. 948 --2.244 14.811 6. 510 15.189 14.748 5. 366 15. 252 --5. 491 6.126 20. 230 --2.406 19. 827 7. 351 20.173 19.770 7. 975 25.151 --5. 873 24.800 6. 705 25. 200 --2. 499 24. 849 --0.121 -- 2. 537 8. 417 30. 125 20.834 7.131 30.166 29. 875 --6. 238 -- 2. 518 34..003 8. 686 35. 097 34.871 7. 414 35.129 --6. 208 39.910 7. 552 40. 090 --2. 436 39. 933 8. 766 40. 067 --2. 266 8. 627 45. 037 --5. 999 44.950 7. 522 45. 050 44.963 --5. 648 --2. 024 49. 992 8. 308 50. (108 49.989 7.344 50.011 --5.19l 55.025 7. 040 54. 975 --1.736 55. O18 7. 843 54. 982 --1. 418 7. 258 59. 958 --4. 654 60.057 6. 624 59. 943 6O. 042 --4. 056 --1.086 6. 566 64. 937 65.085 6.106 64. 915 65. 003 --3. 416 5.490 69.892 --. 760 70. 079 5. 782 69. 921 70.108 --. 460 4. 926 74. 907 --2. 766 75.126 4. 780 74. 874 75. 093 --.2'29 4. 017 79. 889 --2.147 80.151 3. 967 79. 849 80. Ill -- 1.507 3. 018 84. 852 --. 132 ;L 039 84. 891 85.148 85.109 --. 076 2. 946 89. 924 -I. O66 90.104 2. 038 89. 896 90, 076 --. 048 1.039 94.961 --.549 95. 053 I. 028 94. 947 95. 039 --, 0,32 .021 100. 000 _. 021 .032 100. 000 100.000 100. 000 L. E. radius: 0.687 L. E. radius: 1.561 T. E. radius: 0.023 T. E. radius: 0.037 Slope of radius through L.E.: 0.190 Slope of radius through L. E.: 0.003

Tests.--The tests of each smooth airfoil section consisted

TABLE VI.--ORDINATES OF NACA 641A212 AIRFOIL

in measurements of the lift, drag, and quarter-chord pitching-

SECTION

moment coefficients at Reynolds numbers of 3 X 106, 6 X 106, [Stations and ordinates given in percent of airfoil chord]

and 9X106. In addition, the lift and drag characteristics

Upper surface Lower surface

of each section were determined at a Reynolds number of

6X 106 with standard roughness applied to the leading edge

Ordinate Station Ordinate Station

of the model. The standard roughness employed on these

0 O 0 0

24-inch-chord models consisted of 0.011-inch-diameter car-

.409 1.013 .591 -.901 -1.075 .648 1.233 .852

borundum grains spread over a surface length of 8 percent

1. 365 -1.338 1.135 1.580 2.365 2._5 2. 635 -1.8o3

of the chord back from the leading edge on the upper _nd

-2. 423 4.849 3.145 5.151 7. 657 -- 2. 874 7.343 3. 846

lower surfaces. The grains were thinly spread to cover from

9. 842 4. 432 10.158 --3.240 --3. 796 14.849 5. 3_ 15.151

5 to 10 percent of this area. In an effort to obtain some

0.060 20.138 --4.200 19.802 24.880 6.584 25.120 -4.482

idea of the effectiveness of the airfoil sections when equipped

--4.000 29.900 6.9_ 80.100 35. 078 --4.741 34. 922 7.1_

with trailing-edge high-lift devices, each section was fitted

7._2 40. 054 -4. 714 39.946 --4. 549 44.970 7.177 45. 030

with a simulated split flap deflected 60 °. Lift measurements

,50.0007 --4.275 49. 903 6. 035 6.5_ 54. 985 --3. 918 55. 015

with the split flap were made at a Reynolds number of

--3. 499 60. 034 6.103 59. 966 64. 900 --3.034 65. 050 5. 544 6 X 10_ with the airfoil leading edge both smooth and rough.

4. 033 69. 930 --2. 537 70. 064 --2.037 75. 075 4.197 74. 925 79. 910 --1.503 80.090 3.403

RESULTS

2.001 84.912 --1.159 85. 088 --. 771 90. 062 1.751 89. 038 94.968 --. 398 95. 032 .888

The results obtained from tests of the seven airfoil sections

.0_ 100.000 --. 025 100.000

are presented in figures 4 to 10 in the form of standard aero-

L. E. radius: 0.994

T. E. radius: 0.028 dynamic coefficients representing the lift, drag, and quarter-

Slope of radius through L. E.: 0.695 chord pitching-moment characteristics of the airfoil sections.

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THEORETICAL AND EXPERIMENTAL DATA FOR A NUMBER OF NACA 6A--SERIES AIRFOIL SECTIONS 13

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The calculated position of the aerodynamic center and the

sections is about the same as that of removing the cust

variation of the pitching-moment coefficient with lift coeffi-

from the NACA 64-series sections.

cient about this point are also included in these data. The

The comparative data showing the effects upon the aero-

influence of the tunnel boundaries has been removed from

dynamic characteristics of removing the trailing-edge cus_

all the aerodynamic data by means of the following equa- from NACA 6-series airfoil sections should be used witl tions (developed in reference 1):

caution if the cusp removal is affected in some manner othel

than that indicated earlier in this paper. For example, il

cd_O.990Cd'

_he cusp should be removed fl'om a cambered airfoil by meam

of a straight-line fairing of the airfoil surfaces, the amount ol

Cl=0.973CL

camber would be decreased near the trailing edge. Naturall:y

the effect upon the aerodynamic characteristics of removin_

cm_:4=0.951c,_/4_

the cusp in such a manner would not be the same as in-

dicated by the comparative results presented for NACA

no: 1.015ao'

6-series and 6A-series airfoils.

Drag.--The variation of section minimum drag coefficien|

where the primed quantities denote the measured coefficients.

with airfoil thickness ratio at a Reynolds number of 6X 10 _

is shown in figure 11 for NACA 64-series and NACA 64A-

DISCUSSION

series airfoil sections of various cambers, both smooth and

Although the amount of systematic aerodynamic data pre-

with standard leading-edge roughness. As with the NACA

sented for NACA 6A-series airfoil sections is not large, it is

64-series sections (reference 1), the minimum drag coeffi-

enough to indicate the relative merits of the NACA 6A-

cients of the NACA 64A-series sections show no consistent

series airfoil sections as compared with the NACA 6-series

variation with camber. Comparison of the data of figure 11

sections. The variation of the important aerodynamic char-

indicates that removing the cusp from the trailing edge has

acteristics of the five NACA 64A-series airfoils with the

no appreciable effect upon the minimum drag coefficients of

pertinent geometrical parameters of the airfoils is shown in

the airfoils, either smooth or with standard leading-edge

figures 11 to 17, together with comparable data for NACA

roughness.

64-series airfoils. The curves shown in figures 11 to 17 are

Increasing the Reynolds number from 3X106 to 9X 106

for the NACA 64-series airfoil sections and are taken from

has about the same effect upon the minimum drag coefficient

the faired data of reference 1. The experimental points

of NACA 64A-series airfoils (figs. 4 to 10) as that indicated

which appear on these figures represent the results obtained

in reference 1 for the NACA 64-series airfoils.

for the NACA 64A-series airfoil sections in the present

Some differences exist in the drag coefficients of NACA

investigation. Since only two NACA 63A-series sections

64- and 64A-series airfoils outside the low-drag range of lift

were tested, comparative results are not presented for them.

coefficients but these differences are small and do not show

The effect of removing the cusp from the NACA 63-series

any consistent trends (figs. 4 to 10 and reference 1).

.012

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--- 13 ,2 NACA 64A-ser?es

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<> .4

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Srnoofh_

NACA 64-series- .Ol 0

..... Rough j

F.

._ .008 _ . 006 1 I ._.

_....._.-----_I I I .004 ._.

.002 0 2 4 6 8 I0 12 14 16 18 BO 2,- ° A/rfo// fh/'Chness_ percent of" chord FIGURE 11.--Variation n[lllinimtlnl section drag eoellicient with airfoil thickness for some N ACA 6.l-series (reference I) and NAC A 64A-series airfoil sections of various etinlhers ill the smooth condition _md with standard leading-edge roughness. R=6X1U6; flagged symbols indicate NACA 64A-series sections with standard roughness.

THEORETICAL AND EXPERIMENTAL DATA FOR A NUMBER OF NACA 6A--SERIES AIRFOIL SECTIONS 19

T,ift.--The section angle of zero lift as a function of thick- balanced by the increase in lift-curve slope with thickness ncss ratio is shown in figure 12 for NACA 64- and 64A-series ratio shown by NACA 6-series sections. The value of the

airfoil sections of various cambers. These results show that

lift-curve slope for smooth NACA 64A-series airfoil sections

the angle of zero lift is nearly independent of thickness and is very close to that predicted from thin airfoil theory (27r

per red|an or 0.110 per degree). Removing the trailing-

is primarily dependent upon the amount of camber for a

particular type of mean line. Theoretical calculations made edge cusp from an airfoil section with standard leading-edge by use of the mean-line data of figm'e 3 and reference 1 roughness causes the lift-curve slope to decrease quite indicate that airfoils with the a=0.8 (modified) mean line rapidly with increasing airfoil thickness ratio.

The variation of the maximum section lift coefficient with

should have angles of zero lift less negative than those with

the a= 1.0 mean line. Actually, the reverse appears to be airfoil thickness ratio and camber at a Reynolds number

of 6X108 is.shown in figure 14 for NACA 64-series and

the case, and this effect is due mainly to the fact that air-

NACA 64A-series airfoil sections with and without standard

foils having the a= 1.0 type of mean line have angles of zero

lift which are only about 74 percent of their theoretical value leading-edge roughness and simulated split flaps deflected (reference 1), and those having the a----0.8 (modified) mean 60 °. A comparison of these data indicates that the char-

acter of the variation of maximum lift coefficient with airfoil

lines have angles of zero lift larger than indicated by theory•

thickness ratio and camber is nearly the same for the NACA

The measured lift-curve slopes corresponding to the NACA

64-series and NACA 64A-series airfoils of various cambers

64-series and NACA 64A-series airfoil sections. The magni-

are presented in figure 13 as a function of airfoil thickness tude of the maximum lift coefficient appears to be slightly ratio. No consistent variation of lift-curve slope with less for the plain NACA 64A-series airfoils and slightly camber or Reynolds number is shown by either type of air- higher for the NACA 64A-series airfoils with split flaps than foil. The increase in traihng-edge angle which accompanies corresponding values for the NACA 64-series airfoils. These

differences are small, however, and for engineering applica-

removal of the cusp would be expected to reduce the lift-

tions the maximum-lift characteristics of NACA 64-series

curve slope by an amount which increases with airfoil thick-

ness ratio (references 3 and 4). Because the present data and 64A-series airfoil sections of comparable thickness and for the NACA 6A-series sections show essentially no varia- design lift coefficient may be considered practically the same.

tion in lift-curve slope with thickness ratio, it appears that

the effect of increasing the trailing-edge angle is about

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NACA 64-series " - (/VACA 0 0 2 r_ ..0 _ 0---- 0:4" _4 .2 ( o -2 .. -_4

_-4

0 4 8 /2 / 6 20 24 A/r-foE fh/c/<ness_ percent of chord - o .4 .... _! - - F[nURE 12.--Variation of section angle of zero lift with airfoil thickn_s ratio and camber for some NACA 64-series (reference 1) and NACA 64A-series airfoil sections. R=6Xll_ 6.

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./4,

..... _--

s, .... "1"--'1 - - -.

./,2 L Q ./6 • 2}NACA 64A-ser,'es ............

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.... Rough J NAC, 64-series-- A/rfoH /h/c/_ness_ per'cent of chord

I I , I I I

"4 .06 0 4 8 / 2 16 20 24 (a) Airfoil with simulated split flap deflected 60".

A/rfoH th/cktless_ percent of chord (b) Plain airfoil• I,'o;_sRt,: 14. -Variation of maxinmm section lift cocllieicnt with airfoil thickness ratio and Fn_uBl_ 13. Variation of lift-curve slopc with airfoil thickness ratio for some NACA 64-scrio_ camber for some NACA 64-series (reference 1) and NACA 64A-serit,s airfoil sections with (reference I) and N A ('A 64A-sl!rics airfoil sections of w_rious camhcrs hoth in the smooth condition and with standard leading-edge roughness. R=6X10"; flagged symbols indicattl and witlmut simulated split flaps and standard roughness. R_6X10_; flagged symbols indicate NACA fi4A-serics airfoils with standard roughness.

NACA 64A-series sections with standard roughness.

REPORT NO. 903--NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

2O

types of camber. Calculations were made according to thes

A comparison of the maximum-lift data for NACA 64A-

methods for airfoils having the a= 1.0 and a=0.8 (modified

series airfoil sections, presented in figures 4 to 10, with

siinilar data for NACA 64-series airfoil sections indicates mean lines by using the theoretical mean-line data presente, that the scale-effect characteristics of tile two types of section in figure 3 and in reference 1. The results of these calcula

tions indicate that the quarter-chord pitching-moment coeffi

are essentially the same for the range of Reynolds number

from 3 X l06 to 9 X 106. cients of the NACA 64A-series airfoil sections having th

a=0.8 (modified) mean line should be only about 87 percen

Pitohing moment--Thin-airfoil theory provides a means

of those for the NACA 64-series airfoil sections with th

for calculating the theoretical quarter-chord pitching-moment

a=l.0 mean line. The experimental relationship betwee:

coefficients of airfoil sections having various amounts and

the quarter-chord pitching-moment coefficient and airfoi

thickness ratio and camber, shown in figure 15, discloses tha

./

I I% I I

the plain NACA 64A-series airfoils have pitching-momen

I I {_li fNACA 64A-ser/es]__

coefficients which are slightly more negative than those fo

o 0 (NACA G4-ser/es)

J

o .2

the plain NACA 64-series airfoils. The increase in th

"1,- _ 0 --0 .4

I

magnitude of the pitching-moment coefficient of NACA 64A

u I "-'-'-2

series airfoils as compared witlt NACA 64-series airfoil

_ "'.4 (D becomes greater when the airfoils are eq uipped with simulate, 0 -./

split flaps deflected 60 °. A comparison of the theoretic_

and measured pitching-moment coefficients is shown in figur

16 for NACA 64-series and 64A-series airfoil sections. Thes

comparative data indicate that the NACA 64A-series section

much more nearly realize their theoretical moment coefficient

than do the 64-series airfoil sections. Similar trends hay

"_ -_3

been shown to result when mean lines such as the a=0.

type are employed with NACA 6-series airfoils (reference 1)

_erodynamic eenter.--The position of the aerodynami

center and the variation of the moment coefficient with lif

-'4o 4 8 /2 /6 20 24

coefficient about this point were calculated from the quarter

Aii,"*foi/ 7'/"t/ck,"'tes$ t percent of c_oi."d

chord pitching-moment'data for each of the seven airfoil

(a) Plain airfoil. ,, (b) Airfoil with simulated split flap deflected 60°,

tested. The variation of the chordwise position of the nero

FIGURE 15,--Variation of section quarter-chord pitching-moment coefficient at zero angle of

dynamic center with airfoil thickness ratio is shown in figur

attack with airfoil thickness ratio and camber for some NACA 64-series (reference 1) and 17 for the NACA 64-series and 64A-series airfoil section., NAOA 64A-series airfoil sections with and without split flaps. R=6X 106; flagged symbo]s Indicate NAC.& 64A.series airfoils with 60 ° simulated split flap, Since the data for the NACA 64-series airfoils showed n consistent variation with camber, the results are represente, ' -./O

by a single faired curve for all cambers. Following this sam

trend, the position of the aerodynamic center for the NAC_

11 /

64A-series airfoils shows no consistent variation with cambe_

NACA 64A410---- /

The data of figures 4 to 10 show that the variations in th

% "-,08 o

Reynolds number have no consistent effect upon the chord

/

wise position of the aerodynamic center.

Perfect fluid theory indicates that the position of th

641-4/2 / NA CA

_ -.06 aerodynamic center should move rearward with increasin

airfoil thickness and the experimental results for the NAC_

/

64-series airfoil sections follow this trend. The data, c

J _]ACA 64A2/0 .... / ,v,4cA e3Az, .......

....... YVACA 64-2/0

d -.04

....... _VACA 641-2)2 I _. .28 I _i' I L__

,v,,cA 64,A2/z---I !i:o

NACA 64aA2/5-- ...IVACA 642-215 f I y-

27 f

--- a .2 _ _CA 64A-se_/es

/

--NACA _4-_er/e_ .....i-I""I _ :02 ._ .26

/

E o 0 /,

]

"-,'/0 0 -.02 -.04 -.06 _ 08 2 4 6 8 /0 12 14 /6 /8 20 2, T/Teore f/col moment coe?_'clenf for o/rfoil A/rfo_7 f/_/ckmess, percent chord meo,,"t //'me obouf querfer-chord pofHf FIGURE 17.--Variation of chordwise position of aerodynamic center with airfoil thicknc ratio for some NACA 64-series (reference 1) and 64A-series airfoil sections of differc_ FIGtT:a'R. 16.--Comparison of theoretical and measured pitching-moment coefficients for some NACA 64-series and 64A-series airfoil sections. R=6XI06. cambers. R=6XI0 _.

TI-IEORETICAL AND EXPERIMENTAL DATA FOR A NUMBER OF NACA 6A--SERIES AIRFOIL SECTIONS

reference 5 show important forward movements of the aero-

3. Tile section angles of zero lift of NACA 6A-seri(

dynamic center with increasing trailing-edge angle for a

airfoil sections are slightly more negative than those (

given airfoil thickness ratio. The results obtained for tile

comparable NACA 6-series airfoil sections.

NACA 24-, 44-, and 230-series airfoil sections (reference 1)

4. The section quarter-chord pitching-moment coeificient

reveal that the effect of increasing trailing-edge angle pre-

of NACA 6A-series airfoil sections are slightly more negativ

dominates over the effect of increasing thickness because the

than those of comparable NACA 6-series airfoil sectiom

position of the aerodynamic center moves forward with

The position of the aerodynamic center is essentially ind¢

increasing thickness ratio for these airfoil sections. For the

pendent of airfoil thickness ratio for NACA 6A-series airfo

NACA 64A-series airfoils (fig. 17) the aerodynamic center is

sections.

slightly behind the quarter-chord point and does not appear

to vary with increasing thickness. These results suggest

that the effect of increasing thickness is counterbalanced by

increasing trailing-edge angle for these airfoil sections.

LANGLEY _EMOR1AL AERONAUTICAL LABORATORY, CONCLUSIONS NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS, LANGLEY FIELD, VA., _l/iay 6, 1957.

From a two dimensional wind-tunnel investigation of the

aerodynamic characteristics of five NACA 64A-series and

REFERENCES

two NACA 63A-series airfoil sections the following conclu-

1. Abbott, Ira H., Von Doenhoff, Albert E., and Stivers, Louis S

sions based upon data obtained at Reynolds numbers of

Jr.: Summary of Airfoil Data. NACA Rep. No. 824, 1945.

3X106, 6X106, and 9X106 may be drawn: 2. Jacobs, Eastman N., Ward, Kenneth E., and Pinkerton, Robert M.

1. The section minimum drag and maximum lift coef-

The Characteristics of 78 Related Airfoil Sections from Tests i_

ficients of corresponding NACA 6-series and 6A-series airfoil

the Variable-Density Wind Tunnel. NACA Rep. No. 460, 1933 sections are essentially the same. 3. Purser, Paul E., and McKee, John W.: Wind-Tunnel InvestigatioJ of a Plain Aileron with Thickened and Beveled Trailing Edges ol

2. The lift-curve slopes of smooth NACA 6A-series airfoil

a Tapered Low-Drag Wing. NACA ACR. Jan. 1943.

sections appear to be essentially independent of airfoil

4. Jones, Robert T., and Ames, Milton B., Jr. : Wind-Tunnel Investiga

thickness ratio, in contrast to the trends shown by NACA

tion of Control-Surface Characteristics. V--The Use of Beveled Tl'ailing Edge to Reduce the Hinge Moment of a Contro

6-series airfoil sections. The addition of standard leading-

Surface. NACA ARR, March 1942.

edge roughness causes the lift-curve slope to decrease with

5. Purser, Paul E., and Johnson, Harold S.: Effects of Trailing-Edg(

increasing airfoil thickness ratio for NACA 6A-series airfoil

Modifications on Pitching-Moment Characteristics of Airfoils sections.

NACA CB No. L4130, 1944.

u. s. GOVERNMENT PRINTING O.:FICE: 1950